EP1418331A2 - Amortisseur de bruit des turbines à gaz - Google Patents

Amortisseur de bruit des turbines à gaz Download PDF

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Publication number
EP1418331A2
EP1418331A2 EP03256548A EP03256548A EP1418331A2 EP 1418331 A2 EP1418331 A2 EP 1418331A2 EP 03256548 A EP03256548 A EP 03256548A EP 03256548 A EP03256548 A EP 03256548A EP 1418331 A2 EP1418331 A2 EP 1418331A2
Authority
EP
European Patent Office
Prior art keywords
fluid
nozzle
exit
nozzle body
body portion
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP03256548A
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German (de)
English (en)
Other versions
EP1418331A3 (fr
Inventor
John Richard Webster
Paul Jonathan Railton Strange
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Publication of EP1418331A2 publication Critical patent/EP1418331A2/fr
Publication of EP1418331A3 publication Critical patent/EP1418331A3/fr
Withdrawn legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/28—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow
    • F02K1/34—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow for attenuating noise

Definitions

  • Embodiments of the present invention relate to the suppression of part of the noise arising from the fluid flow output by a gas turbine engine. They particularly relate to the suppression of those parts of the noise which are audible to humans.
  • bypass nozzle fed by a core bypass reduces the mean velocity of the engine's exhaust products and reduces the noise of the engine. This is, however, a partial solution. There are limits to the size of the bypass because as the bypass size is increased, although the mean velocity continues to drop, the engine size and drag increases.
  • a current additional solution to the problem of noise is to use forced mixers at the exit of the hot nozzle.
  • a disadvantage of this is that it requires the bypass to extend beyond the exhaust of the core nozzle, which increases the weight of the engine.
  • US2990905 discloses an apparatus for suppressing part of the noise created in jet engines.
  • the core nozzle ending comprises co-axial inner and outer walls separated by an annular duct.
  • the annular duct is connected to a helium supply or an air supply tapped from the core and has a series of nozzles in the downstream wall of the annular duct.
  • These nozzles emit auxiliary jets which penetrate an outer envelope of the main jet.
  • the auxiliary jets set the pattern of the turbulence in the main jet and are in some ways analogous to the effect of teeth or corrugations in the main jet flow.
  • a problem with this solution is that the flat end to the core nozzle created by the annular duct is not aerodynamically efficient and produces drag. There is therefore a permanent performance penalty in comparison to a tapered core nozzle ending which terminates at a knife-edge.
  • a gas turbine engine (10) comprising: a nozzle (19, 21) comprising a nozzle body portion (32, 33, 74) defining a nozzle exit (42, 43), characterised in that the nozzle body portion (32, 33, 74) comprises fluid injection means (62), positioned upstream of the exit (42, 43) relative to a fluid flow (F1, F2) created by the operation of the engine, for injecting fluid (68) upstream of the exit (42, 43).
  • a method of suppressing part of the noise of a gas turbine engine comprising a nozzle (19, 21), the nozzle (19, 21) comprising a nozzle body portion (32, 33, 74) defining a nozzle exit (42, 43), the method comprising the step of: injecting fluid (68) into a fluid flow (F1) created by the operation of the engine (10) while the fluid flow (F1) is travelling adjacent the nozzle body portion (32, 33, 74).
  • a gas turbine engine (10) comprising: a nozzle (19, 21) comprising a nozzle body portion (32, 33, 74) defining a nozzle exit (42, 43), characterised in that the nozzle body portion (32, 33, 74) comprises output means (62), positioned upstream of the exit (42, 43) relative to a fluid flow (F1) created by the operation of the engine (10), for disturbing a boundary layer (75) between the nozzle body portion (32, 33, 74) and the fluid flow (F1).
  • a method of suppressing part of the noise of a gas turbine engine comprising a nozzle (19, 21), the nozzle (19, 21) comprising a nozzle body portion (32, 33) defining a nozzle exit (42, 43), the method comprising the step of: disturbing a boundary layer between the nozzle body portion and a fluid flow created by the operation of the engine.
  • the boundary layer is disturbed as opposed to disturbing the shear layer using air jets or tabs/serrations, as in the prior art. A smaller amount of energy is required to manipulate the boundary layer compared to the shear layer, and achieve comparable suppression of noise.
  • a jet engine (10) for an aeroplane comprising: a nozzle (19, 21), the nozzle comprising a nozzle body portion (32, 33) comprising fluid injection means for injecting fluid characterised in that the jet engine further comprises control means for controlling the fluid injection means to inject fluid during a first phase of operation and to not inject fluid during a second phase of operation.
  • Fig. 1 illustrates a sectional side view of the upper half of a gas turbine engine 10.
  • the gas turbine engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18 and a core nozzle 19.
  • the compressors 13, 14, 15, the combustor 15 and the turbine arrangement form the core of the engine.
  • the gas turbine engine 10 has core bypass 20 connected between the propulsive fan 12 and a bypass nozzle 21, inscribing the hot exhaust nozzle 19.
  • the gas turbine engine 10 operates in a conventional manner so that air entering in the intake 11 is accelerated by the propulsive fan 12 which produces two air flows: a first air flow into the core and a second air flow into the by-pass 20 which provides propulsive thrust.
  • the intermediate pressure compressor 13 compresses the air flow directed into it for delivery to the high pressure compressor 14 where further compression takes place.
  • the compressed air from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted.
  • the resultant hot combustion products then expand and thereby drive the high, intermediate and low pressure turbines 16, 17, 18 before being exhausted through the core nozzle 19 to provide additional propulsive thrust.
  • the high, intermediate and low pressure turbines 16, 17, 18 respectively drive the high and intermediate pressure compressors 14, 13 and the propulsive fan 12 by suitable interconnecting shafts 22.
  • the direction of fluid flow in the figure is therefore from left to right.
  • the bypass nozzle 21 is annular and defines a bypass nozzle flow channel 30.
  • the bypass nozzle flow channel 30 is bounded on its outside edge by an interior surface 38b of a bypass nozzle body portion 32 and on its inside edge by an exterior surface 39a of a core nozzle body portion 33. In other embodiments by bypass nozzle 21 may be circular.
  • the core nozzle 19 is annular and defines a core nozzle flow channel 31.
  • the core nozzle flow channel 31 is bounded on its outside edge by an interior surface 39b of the core nozzle body potion 33 and on its inside edge by an exterior surface 40 of a plug 34.
  • the core nozzle may be circular.
  • the bypass nozzle body portion 32 is part of an outer body portion 36 of the gas turbine engine 10. It has an exterior surface 38a and an interior surface 38b that converge to meet at an acute angle at the exit 42 of the bypass nozzle 21.
  • the bypass nozzle body portion 32 therefore tapers at its ending to an edge at the exit 42 of the bypass nozzle 21.
  • the core nozzle body portion 33 is part of an inner body portion 37 of the gas turbine engine 10. It has an exterior surface 39a and an interior surface 39b that converge to meet at an acute angle at the exit 43 of the core nozzle 19. The core nozzle body portion 33 therefore tapers at its ending to an edge at the exit 43 of the core nozzle 43.
  • Embodiments of the invention cause an aerodynamic disturbance upstream of a nozzle exit in order to enhance mixing and hence reduce noise.
  • the aerodynamic disturbance is preferably formed adjacent a nozzle body portion. For example it may be created at one or more locations adjacent any one or more of the exterior surface 38a of the bypass nozzle body 32, the interior surface 38b of the bypass nozzle body 32, the exterior surface 39a of the core nozzle body 33 and the interior surface 39b of the core nozzle body 33.
  • the aerodynamic disturbance is created by outputting energy into a fluid stream using for example sound wave production means or, preferably, fluid injection means.
  • Figure 2 illustrates a nozzle body 74.
  • the nozzle body 74 has a first surface 70 which may be the upper surface or lower surface of the core nozzle body or the upper surface or the lower surface of the bypass nozzle body.
  • the nozzle body 74 has a second surface 72 which would be respectively the lower or upper surface of the core nozzle body or the lower or upper surface of the bypass nozzle body.
  • a first fluid flow F1 flows over first surface 70 of the nozzle body 74.
  • a second fluid flow flows over the second surface 72.
  • the nozzle body 74 is the bypass nozzle body, then one of the fluid flows F1, F2 is produced by the fluid exhausted from the bypass and the other is produced, for example, by the propulsion of the nozzle body through the atmosphere as part of an aeroplane engine and has a lower speed. If the nozzle body 74 is a core nozzle body, then one of the fluid flows F1, F2 is produced by the fluid exhausted from the core and the other is produced by the fluid exhausted by the bypass.
  • the nozzle body 74 has fluid injection means 60 comprising an aperture 62 in the first surface 70 of the nozzle body 74.
  • the aperture 60 is connected via a feed 64 to a supply of air 66.
  • the air supplied exits the aperture 62 as a air jet 68, which forms an acute angle ⁇ with the first surface 70 of the nozzle.
  • the air jet will preferably have an axial component of velocity in the direction of the fluid flow F1 and a radial component of velocity into the fluid flow F1, but it may also have a tangential component of velocity (into or out of the page of the figure).
  • the air jet 68 is preferably directed downstream, that is in the direction of the fluid flow F1, and enters the fluid flow F1 adjacent the aperture 62.
  • the fluid flow F1 has a boundary layer 75 of static or very slow moving air adjacent the first surface 70 upstream of the aperture 62.
  • the air jet 68 disturbs the boundary layer 75 at the aperture 62 and downstream of the aperture 62.
  • the boundary layer 75 grows creating a low speed turbulent region 76 adjacent the first surface 70 downstream of the aperture which increases in size as it progresses down stream to the nozzle exit.
  • the low speed turbulent region 76 causes the fluid flow F2 to turn into the fluid flow F1 and increases the rate of mixing of the fluid flows F1 and F2 in the shear layer 77.
  • a nozzle body may have a plurality of apertures on one of its surfaces. Each aperture may produce an air jet 68 which is at the same angle and direction (axial, radial and tangential components) or each may have a different angle and direction. There may be apertures in one, all or any combination of the exterior surface 38a of the bypass nozzle body 32, the interior surface 38b of the bypass nozzle body 32, the exterior surface 39a of the core nozzle body 33 and the interior surface 39b of the core nozzle body 33.
  • the apertures used may be particularly small e.g. less than a few mms, in which case the air jets 68 are called microjets.
  • the mass flow through the apertures may be around 1% of the core air flow.
  • the use of microjets, and in particular, pulsed microjets allows a lower mass flow to be used in the air jet 68 to achieve the desired boundary layer disturbance. Pulsed microjets are particularly useful for achieving boundary layer separation.
  • Figure 3a illustrates an air supply control mechanism 80 comprising a tap 82 from one of the compressors of the core which supplies pressurised air, a switchable valve mechanism 84 and an output 86 for providing the air supply 66.
  • An input control signal 88 is used to control the switchable valve mechanism 84 between an open position in which the tap 82 is connected to the output 86 and a closed position in which the connection between tap 82 and output 86 is closed.
  • the output 86 may connect directly to an air jet feed 64 or via a manifold to a plurality of air jet feeds.
  • the switchable valve is arranged to control the mass flow of air in an air jet 68 as well as switching the air jet 68 on and off.
  • Figure 3b illustrates a microjet air supply mechanism 90 comprising a tap 82 from one of the compressors of the core, a switchable valve mechanism 84 connected to the tap 82, a pulsing mechanism 92 connected to the switchable valve mechanism 84 and an output 86 connected to the pulsing mechanism 92, for providing the air supply 66.
  • An input control signal 88 is used to control the switchable valve mechanism 84 between an open position in which the tap 82 is connected to the pulsing mechanism 92 and a closed position in which the connection between tap 82 and pulsing mechanism 92 is closed.
  • the switchable valve mechanism 84 is arranged to control the mass flow of air in an air jet 68 (i.e.
  • the pulsing mechanism 92 receives a second input control signal 94 that controls the frequency at which the air supply 66 is pulsed.
  • the pulsing frequency is fixed and the pulsing mechanism 92 may be a simple pneumatic or mechanical oscillator which is activated when it receives an air flow (e.g. a whistle or horn).
  • the frequency of the oscillator is preferably tuned to the natural length of the boundary layer which is of the order of kHz.
  • the output 86 of the pulsing mechanism 92 may connect directly to an air jet feed 64 or via a manifold to a plurality of air jet feeds.
  • the output of the switchable valve mechanism 84 may connect only to a single pulsing mechanism 92 or via a manifold to a plurality of pulsing mechanisms 92, each pulsing mechanism 92 being capable of providing one or more pulsed air jets.
  • the switchable valve mechanism 84 may be used to turn the air jets on and off. This may occur regularly with a frequency of Hz in order to obtain a desired disruption of the boundary layer 80.
  • the switchable valve mechanism can be used to switch on the air jets and reduce noise at take-off when the aeroplane engine is particularly noisy and close to the ground. The mechanism can then be used to switch off the air jets when the aeroplane is cruising at altitude.
  • the air jets can be used to suppress part of the jet engine noise at take-off and can be switched off when noise suppression is no longer required to achieve maximum fuel efficiency.
  • Figure 4 illustrates an end view into a nozzle of a gas turbine engine 10.
  • the edge 42 of the bypass nozzle body 32 , the edge 43 of the core nozzle body 33 and the plug 34 are illustrated.
  • Also illustrated are a series of air jets 68a injected from the exterior surface 39a of the core nozzle body 33 into the fluid flow F2 and a series of air jets 68b injected from the interior surface 39b of the core nozzle body 33 into the fluid flow F1.
  • the series of air jets 68a are injected at an angle ⁇ relative to the tangent to the exterior surface 39a. In this example all the air jets 68a are injected at the same angle which is approximately 90 degrees i.e.
  • the angle ⁇ of the air jets may be different for each air jet 68a.
  • the series of air jets 68b are injected at an angle ⁇ relative to the tangent to the interior surface 39b. In this example all the air jets 68b are injected at the same angle which is approximately 90 degrees i.e. with no tangential component, but in other embodiments the angle ⁇ of the air jets may be different for each air jet 68b.
  • each of the series of air jets 68/68b may have their own supply control mechanism 80/90 in which case the amplitude and/or frequency of each air jet can be separately controlled, or a plurality of the air jets may be feed from a manifold connected to a supply control mechanism, in which case the amplitude and/or frequency of the plurality of air jets can be controlled together.
  • the air jets 68a and 68b are operating simultaneously.
  • the air jets 68a are separately controlled to the air jets 68b and the operation of the air jets 68a and 68b alternates.
  • the air jets generally have a fixed position and a fixed angle. However, they may be switched on or off in unison, in groups or individually.
  • the mass flow of an air jet may be fixed with all of the air jets having the same mass flow or different air jets having different mass flow. Alternatively, the mass flow of an air jet may be altered. Such alteration may occur in unison, in groups or individually.
  • the air jets may be pulsed. The pulsing may be at a fixed or variable frequency. The pulsing may be applied selectively to some or all of the air jets. It is therefore possible to modulate the disturbance created by the air jets.
  • the core nozzle and bypass nozzle exit planes are substantially co-planar the invention is applicable to different nozzle geometries, such as when the core nozzle is recessed within the bypass nozzle for internal mixing, when the core nozzle extends beyond the bypass nozzle for external mixing or when no bypass is used and gas is exhausted via a single core nozzle.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Nozzles (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Control Of Turbines (AREA)
  • Jet Pumps And Other Pumps (AREA)
EP03256548A 2002-11-09 2003-10-17 Amortisseur de bruit des turbines à gaz Withdrawn EP1418331A3 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GBGB0226228.5A GB0226228D0 (en) 2002-11-09 2002-11-09 Suppression of part of the noise from a gas turbine engine
GB0226228 2002-11-09

Publications (2)

Publication Number Publication Date
EP1418331A2 true EP1418331A2 (fr) 2004-05-12
EP1418331A3 EP1418331A3 (fr) 2006-07-12

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EP03256548A Withdrawn EP1418331A3 (fr) 2002-11-09 2003-10-17 Amortisseur de bruit des turbines à gaz

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US (2) US20040088967A1 (fr)
EP (1) EP1418331A3 (fr)
GB (1) GB0226228D0 (fr)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2005023929A (ja) * 2003-06-30 2005-01-27 General Electric Co <Ge> ジェット騒音低減のための流体シェブロン及び構成可能な熱シールド
EP1553281A1 (fr) * 2003-12-30 2005-07-13 General Electric Company Dispositif de réduction des bruits d'échappement de turboréacteurs utilisant des jets oscillants
EP1698774A2 (fr) 2005-03-02 2006-09-06 Rolls-Royce Plc L'entraînement d'accessoires dans une période de poussée limitée
FR2901321A1 (fr) * 2006-05-18 2007-11-23 Aircelle Sa Procede d'homogeneisation de l'air en sortie de turboreacteur pour abaisser le bruit genere
EP2256327A1 (fr) * 2008-02-25 2010-12-01 IHI Corporation Dispositif de réduction de bruit et système de propulsion par réaction
DE102021209284A1 (de) 2021-08-24 2023-03-02 Deutsches Zentrum für Luft- und Raumfahrt e.V. Vorrichtung und Verfahren zur Beeinflussung einer Fluidhauptströmung und Düse

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US7055329B2 (en) * 2003-03-31 2006-06-06 General Electric Company Method and apparatus for noise attenuation for gas turbine engines using at least one synthetic jet actuator for injecting air
US7481038B2 (en) * 2004-10-28 2009-01-27 United Technologies Corporation Yaw vectoring for exhaust nozzle
US20060202082A1 (en) * 2005-01-21 2006-09-14 Alvi Farrukh S Microjet actuators for the control of flow separation and distortion
US20090261206A1 (en) * 2005-01-21 2009-10-22 Alvi Farrukh S Method of using microjet actuators for the control of flow separation and distortion
US7793504B2 (en) * 2006-05-04 2010-09-14 Rolls-Royce Corporation Nozzle with an adjustable throat
US8001762B2 (en) * 2006-08-04 2011-08-23 Efremkin Artem P Method and device to increase thrust and efficiency of jet engine
US7966824B2 (en) 2006-08-09 2011-06-28 The Boeing Company Jet engine nozzle exit configurations and associated systems and methods
US8157207B2 (en) * 2006-08-09 2012-04-17 The Boeing Company Jet engine nozzle exit configurations, including projections oriented relative to pylons, and associated systems and methods
US8015819B2 (en) * 2006-09-29 2011-09-13 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Wet active chevron nozzle for controllable jet noise reduction
US7870722B2 (en) * 2006-12-06 2011-01-18 The Boeing Company Systems and methods for passively directing aircraft engine nozzle flows
US7966826B2 (en) * 2007-02-14 2011-06-28 The Boeing Company Systems and methods for reducing noise from jet engine exhaust
FR2920035B1 (fr) * 2007-08-17 2013-09-06 Airbus France Turbomoteur a emission de bruit reduite pour aeronef
US9528468B2 (en) * 2009-10-28 2016-12-27 Ihi Corporation Noise reduction system
FR2960028B1 (fr) * 2010-05-12 2016-07-15 Snecma Dispositif pour attenuer le bruit emis par le jet d'un moteur de propulsion d'un aeronef.
US8316631B2 (en) * 2010-09-30 2012-11-27 Lockheed Martin Corporation Exhaust plume heat effect reducing method and apparatus
FR2981134B1 (fr) * 2011-10-06 2014-09-19 Snecma Dispositif avec paroi avec au moins deux ouvertures debouchant dans un flux de gaz
US20130305731A1 (en) * 2012-05-17 2013-11-21 Philip John MORRIS Methods and apparatus for providing fluidic inserts into an exhaust stream to reduce jet noise from a nozzle
US9970386B2 (en) * 2013-06-07 2018-05-15 United Technologies Corporation Exhaust stream mixer
US9541030B2 (en) * 2013-11-27 2017-01-10 Lockheed Martin Corporation Exhaust plume cooling using periodic interruption of exhaust gas flow to form ambient air entraining vortices
US11365704B2 (en) * 2018-02-27 2022-06-21 New York University In Abu Dhabi Corportion Directionally targeted jet noise reduction system and method

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Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2005023929A (ja) * 2003-06-30 2005-01-27 General Electric Co <Ge> ジェット騒音低減のための流体シェブロン及び構成可能な熱シールド
EP1553281A1 (fr) * 2003-12-30 2005-07-13 General Electric Company Dispositif de réduction des bruits d'échappement de turboréacteurs utilisant des jets oscillants
US7308966B2 (en) 2003-12-30 2007-12-18 General Electric Company Device for reducing jet engine exhaust noise using oscillating jets
EP1698774A2 (fr) 2005-03-02 2006-09-06 Rolls-Royce Plc L'entraînement d'accessoires dans une période de poussée limitée
EP1698774A3 (fr) * 2005-03-02 2012-11-21 Rolls-Royce Plc L'entraînement d'accessoires dans une période de poussée limitée
FR2901321A1 (fr) * 2006-05-18 2007-11-23 Aircelle Sa Procede d'homogeneisation de l'air en sortie de turboreacteur pour abaisser le bruit genere
WO2007135257A1 (fr) * 2006-05-18 2007-11-29 Aircelle Nacelle de turboreacteur equipee de moyens de reduction du bruit engendre par ce turboreacteur
RU2445489C2 (ru) * 2006-05-18 2012-03-20 Эрсель Гондола турбореактивного двигателя, снабженная средствами снижения шума, создаваемого этим двигателем
EP2256327A1 (fr) * 2008-02-25 2010-12-01 IHI Corporation Dispositif de réduction de bruit et système de propulsion par réaction
US8904795B2 (en) 2008-02-25 2014-12-09 Ihi Corporation Noise reducing device and jet propulsion system
EP2256327B1 (fr) * 2008-02-25 2019-09-04 IHI Corporation Dispositif de réduction de bruit et système de propulsion par réaction
DE102021209284A1 (de) 2021-08-24 2023-03-02 Deutsches Zentrum für Luft- und Raumfahrt e.V. Vorrichtung und Verfahren zur Beeinflussung einer Fluidhauptströmung und Düse

Also Published As

Publication number Publication date
GB0226228D0 (en) 2002-12-18
EP1418331A3 (fr) 2006-07-12
US20040088967A1 (en) 2004-05-13
US20060283188A1 (en) 2006-12-21

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